Geared Turbine Engine with Relatively Lightweight Propulsor Module

ABSTRACT

An example gas turbine engine includes a propulsor assembly consisting of a propulsor and a first turbine. An epicyclic gear train defines a gear reduction ratio of greater than 2.3. A weight of the propulsor assembly is less than 40 percent of a total weight of the gas turbine engine, excluding any nacelle. A first spool includes a first shaft that connects a first compressor and the first turbine. The first shaft drives the propulsor through the gear train to drive the propulsor at a lower speed than the first spool. A second spool includes a second shaft connecting a second compressor and a second turbine. A weight of the propulsor is greater than a weight of the first turbine.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/555,066 filed Dec. 20, 2021, which is a continuation of U.S. patentapplication Ser. No. 16/999,507 filed Aug. 21, 2020, which is acontinuation of U.S. patent application Ser. No. 16/152,710 filed Oct.5, 2018, which is a continuation of U.S. patent application No.14/432,377 filed Mar. 30, 2015, now U.S. Pat. No. 10,100,745 grantedOct. 16, 2018, which is a national stage entry of InternationalApplication No. PCT/US2013/025276 filed Feb. 8, 2013, and which claimspriority to United States Provisional Application No. 61/710,808 filedon 8 Oct. 2012, and is incorporated herein by reference.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The high pressure turbine drives the high pressure compressor through anouter shaft to form a high spool, and the low pressure turbine drivesthe low pressure compressor through an inner shaft to form a low spool.The fan section may also be driven by the low inner shaft. A speedreduction device such as an epicyclical gear assembly may be utilized todrive the fan section such that the fan section may rotate at a speeddifferent than the turbine section so as to increase the overallpropulsive efficiency of the engine. In such engine architectures, ashaft driven by one of the turbine sections provides an input to theepicyclical gear assembly that drives the fan section at a reduced speedsuch that both the turbine section and the fan section can rotate atcloser to optimal speeds.

Although geared architectures have improved propulsive efficiency,turbine engine manufacturers continue to seek further improvements toengine performance including improvements to thermal, transfer, andpropulsive efficiencies.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes, among other things, a propulsor assembly includingat least a fan module and a fan drive turbine module; a gas generatorassembly including at least a compressor section, a combustor in fluidcommunication with the compressor section; and a turbine in fluidcommunication with the combustor; and a geared architecture driven bythe fan drive turbine module for rotating a fan of the fan module. Aweight of the fan module and the fan drive turbine module is less thanabout 40% of a total weight of a gas turbine engine.

In a further non-limiting embodiment of the foregoing gas turbineengine, the fan module comprises no more than 26 fan blades.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the fan module comprises more than 26 shrouded fanblades.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises a rotor that isconfigured to rotate more than 2.6 times for every single rotation ofthe fan.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises directionally solidifiedblades.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises fewer than six stages.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the weight of the fan module and the fan drive turbine moduleis from 28 to 34 percent the total weight of the gas turbine engine.

A gas turbine engine according to another exemplary aspect of thepresent disclosure includes, among other things, a propulsor assembly ofa gas turbine engine, the propulsor assembly including at least a fanmodule and a fan drive turbine module, the propulsor assembly is lessthan about 40% of a total weight of a gas turbine engine.

In a further non-limiting embodiment of the foregoing gas turbineengines, the fan module comprises no more than 26 fan blades.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the fan module comprises more than 26 shrouded fanblades.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises a rotor that isconfigured to rotate 2.6 times for every single rotation of the fan.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises directionally solidifiedblades.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan drive turbine module comprises fewer than six stages.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the propulsor assembly is from 28 to 34 percent the totalweight of the gas turbine engine.

A method of distributing weight between a propulsor assembly and a gasgenerator assembly of a gas turbine engine according to anotherexemplary aspect of the present disclosure includes, among other things,providing a propulsor assembly that have a first weight, the propulsorassembly including a fan module and a turbine module; and configuringthe propulsor assembly for installation within a gas turbine enginehaving a second weight when the propulsor assembly is installed, whereinthe first weight is less than 40 percent of the second weight.

In a further non-limiting embodiment of the foregoing method ofdistributing weight, the fan module comprises no more than 26 fanblades.

In a further non-limiting embodiment of either of the foregoing methodsof distributing weight, the fan module comprises more than 26 shroudedfan blades.

In a further non-limiting embodiment of any of the foregoing methods ofdistributing weight, the fan drive turbine module comprises a rotor thatis configured to rotate 2.6 times for every single rotation of the fan.

In a further non-limiting embodiment of any of the foregoing methods ofdistributing weight, the propulsor assembly is from 28 and 34 percent atotal weight of a gas turbine engine.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the detaileddescription. The figures that accompany the detailed description can bebriefly described as follows:

FIG. 1 shows a section view of an example gas turbine engine.

FIG. 2 shows a section view of a portion of an example embodiment of thegas turbine engine of FIG. 1 .

FIG. 3 shows a comparative table of features of the gas turbine engineof FIG. 2 and other gas turbine engines.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a gas turbine gasturbine engine, it should be understood that the concepts describedherein are not limited to use with gas turbines as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44 thenby the high pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expandedthrough the high pressure turbine 54 and low pressure turbine 46. Themid-turbine frame 58 includes vanes 60, which are in the core airflowpath and function as an inlet guide vane for the low pressure turbine46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guidevane for low pressure turbine 46 decreases the length of the lowpressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of the turbine section28. Thus, the compactness of the gas turbine engine 20 is increased anda higher power density may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as bucket cruiseThrust Specific Fuel Consumption (TSFC)—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7° R)]{circumflex over ( )}0.5. The “Low corrected fan tipspeed”, as disclosed herein according to one non-limiting embodiment, isless than about 1150 ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 20fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 34. In another non-limiting example embodiment, the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan blades and the number of low pressure turbine rotors isbetween about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 34 in the lowpressure turbine 46 and the number of blades in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

The example gas turbine engine 20 includes weight reduction featuresfacilitating improved efficiency. Example weight reduction featuresprovide a propulsor assembly in the engine 20 that, in total, is lessthan about 40% of the total engine weight. Engines having a propulsorassembly that is less than about 40% of the total engine weight havebeen found to have a more efficient and targeted weight distributionthan other engines.

Engines having weight distributed in this way have relatively lighterfront ends, which may be advantageous as the engine 20 is cantileveredforward of the wing. For example, a pylon structure (not shown) securingthe engine 20 to a wing must typically hold the engine 20 under veryhigh g loads and even crash loads. The greater the weight of the fansection, the greater the weight of the pylon structure. The engine 20and pylon structure are held by the wing where both the weight of theengine 20 and the moment arm of the fan section 22 and the low pressureturbine 46 and the pylon structure must be accommodated.

Referring to FIGS. 2 and 3 with continued reference to FIG. 1 , anexample gas turbine engine 20 a includes a propulsor assembly 62 and agas generator assembly 64. In this example, the propulsor assembly 62includes a fan module 66 and a turbine module 68. Generally, thepropulsor assembly 62 includes structures associated with producingthrust. The gas generator assembly 64 includes the remaining portions ofthe engine 20 a. In this example, the turbine module 68 is a lowpressure, or fan drive, turbine module.

As known, modular construction of gas turbine engines has developed tofacilitate assembly, transportation, and repair. A person having skillin this art in the benefit of this disclosure would understand thegeneral boundaries of the propulsor assembly 62 and gas generatorassembly 64 within a gas turbine engine 20 a, as well as the modulestherein.

The example fan module 66 includes the fan 42. The fan 42 includes a hub70 and an array of blades 72 extending radially from the hub 70. The hub70 and blades 72 fit within an annular fan case 76.

A nacelle 74 circumscribes the fan module 66 and other portions of theengine 20 a. In this example, a front flange 78 and a rear flange 82 areused to secure the fan module 66 to the nacelle 74 and the gas turbineengine 20 a. The terms front and rear are with reference to a generaldirection of flow through the engine 20 a.

In this example, the front flange 78 directly secures the case 76 of thefan module 66 to the nacelle 74 at a position axially forward the blades72. Components axially forward and radially outward of the flange 78 areconsidered portions of the nacelle 74.

In this example, the rear flange 82 directly secures the case 76 of thefan module 66 to the nacelle 74 at a position axially rearward theblades 72. Components axially rearward and radially outboard of the rearflange 82 are considered portions of the nacelle 74. The rear flange 82may attach at a position that is rearward of a fan exit guide vane 86.

In this example, the nacelle 74 is a considered a separate structurefrom the engine 20 a, a thrust reverser system 75, and flanges 78 and82.

The geared architecture 48 of the gas turbine engine 20 a has a bearingcompartment front wall 90. The example fan module 66 includes thebearing compartment front wall 90, but does not include other portionsof the geared architecture 48. The bearing compartment front wall 90supports the fan 42. The bearing compartment front wall 90 is typicallyshipped together with the remaining portions of the fan module 66.

The fan module 66 has a weight F_(W). The nacelle 74, the front flange78, and the rear flange 78 are, in this example, excluded whendetermining the overall weight of the fan module 66.

The turbine module 68 is secured within the engine 20 a by at least afront flange 92, a rear flange 94, and hub bolts 96. The front flange 92secures the turbine module 68 to the mid-turbine frame 58. The rearflange 94 secures the turbine module 68 to a turbine exhaust case 100.The hub bolts 96 secure the turbine module 68 to the inner shaft 40 ofthe low speed spool 30.

The turbine module 68 has a weight T_(W). The mid-turbine frame 58, theexhaust case 100, and the shaft 40 are, in this example, excluded whendetermining the overall weight of the fan module 66.

The propulsor assembly 62 has a total weight P_(TOT), which is the sumof the weight F_(W) of the fan module 66 and the weight T_(W) of theturbine module 68. That is, P_(TOT)=F_(W)+T_(W).

In addition to the propulsor assembly 62, the example engine 20 aincludes a gas generator assembly 64. The structures of the gasgenerator assembly 64 are generally considered to be the portions of theengine 20 a that are not part of the propulsor assembly 62. The gasgenerator assembly 64 has a total weight G_(TOT).

The gas generator assembly 64 thus includes the low pressure compressor44, the high pressure compressor 52, a diffuser case, and the highpressure turbine 54. The gas generator assembly 64 further includes themid turbine frame 58, all bearing systems 38, the inner shaft 40, atower shaft 80, external components, such as an accessory gearbox 88,control and wire harnesses, and pressure sensing devices and tubes, andall other externals and fluids.

In another geared gas turbine configuration utilizing three spools, thegas generator assembly 64 may additionally include an intermediatepressure compressor and intermediate pressure turbine.

As can be appreciated, the engine 20 a has a total weight Eng_(TOT),which can be determined by adding the weight P_(TOT) of the propulsorassembly 62 and the weight G_(TOT) of the gas generator assembly 64.That is, Eng_(TOT)=P_(TOT)+G_(TOT).

Components of the example propulsor assembly 62 include featuresfacilitating reduced the weight P_(TOT) of the propulsor assembly 62. Inthis example, the weight P_(TOT) is less than about 40% of the totalengine weight Eng_(TOT).

Example weight reducing features of the fan module 66 can includeconstructing the blades 72 of one or more relatively lightweightmaterials, such as aluminum, hollow aluminum, hollow titanium, compositematerials and plastic, or some combination of these. The number ofblades 72 in the engine 20 a is less than about 26, which alsocontributes to reducing weight.

In some examples, the fan module 66 may include blades 72 of a shroudedfan blade configuration having more than 26 blades. The blades 72 caninclude a lightweight fan blade leading edge protection featuresincluding, but not limited to, a titanium shroud, nickel shroud, and/ora metallic coating in a leading edge region.

The fan module 66 may further include and be enabled by using alightweight fan blade containment system. A disclosed examplelightweight fan blade containment system could include one of or acombination of aluminum, and/or an organic matrix composite material.

The weight of the fan module 66 can be influenced by how many blades 72are used (few blades 72 may be heavier and more difficult to contain),whether the blades 72 are solid or hollow; whether the blades 72 have anumber greater than 26 and therefore require a shroud between blades 72.The fan blades 72 can be titanium solid (inexpensive, heavy); titaniumhollow (expensive, light); composite with a metal leading edge (light,expensive); solid aluminum (light, inexpensive) or hollow aluminum(ultra-light, inexpensive). The fan case can be aluminum with a Kevlarcontainment system (this is cheap and heavy) or can be a wound compositecase (more expensive, light in weight).

Additional features facilitating use of a relatively lightweightpropulsor assembly 62 within the engine 20 a include tapered rollerbearings that reduce engine length. A canted fan exit guide vane furtherprovides an efficient load connection between the fan rotor support andan outer barrel of the fan section 22.

Example weight reducing features of the turbine module 68 include arelatively high speed low pressure turbine rotor 102 configured tooperate at a rotational speed that is at least 2.6 times the speed ofthe fan 42. That is, the rotor 102 is configured to rotate 2.6 times forevery single rotation of the fan 42.

In this example, the low pressure turbine 46 is a fan drive turbine.Similarly, the turbine module 68 is a fan drive turbine module. Othergeared gas turbine configurations that utilize three turbines may alsoinclude a fan drive turbine operating within similar speed ratio ranges.

In this example, the low pressure turbine 46 includes fewer than aboutsix stages. The number of stages of the low pressure turbine 46 is anexample of many elements that facilitate maintaining the disclosedweight ratio of the propulsor assembly 62 relative to the overall engineweight. Portions of the propulsor assembly, such as the low pressureturbine 46, may include directionally solidified blades.

The low pressure turbine could also be three stages or four. The fourstage version may be more efficient, but heavier, than the three stageversion. The bearing compartment 38 supporting the low pressure turbine46 shaft can be at the far end of the shaft 40 (which may be heavier andless expensive) or between the high pressure turbine 24 and low pressureturbine 46 (which may be lighter, more expensive, hotter and a challengeto design and repair).

In the disclosed example listed in FIG. 3 , the overall engine weightEng_(TOT) (which does not include the nacelle structure and mounts) isabout 6162 lbs (2795 kg) with a propulsor assembly weight P_(TOT) ofabout 1838 lbs (834 kg). The propulsor assembly weight P_(TOT) is about29.8 percent of the total geared gas turbine weight Eng_(TOT) In anotherdisclosed example, the total engine weight Eng_(TOT) is about 4837 lbs(2194 kg) and the propulsor module weight P_(TOT) is about 1604 lbs (728kg) or about 33.2 percent of the total engine weight Eng_(TOT) In afurther disclosed example, the total engine weight Eng_(TOT) is about3637 lbs (1650 kg) and the propulsor module weight P_(TOT) is about 1033(469 kg) or about 28.4 percent of the total engine weight Eng_(TOT).

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

We claim:
 1. A gas turbine engine comprising: a propulsor assemblyconsisting of a propulsor and a first turbine, the propulsor including ahub and an array of no more than 26 blades extending from the hub; anepicyclic gear train defining a gear reduction ratio of greater than2.3; wherein a weight of the propulsor assembly is less than 40 percentof a total weight of the gas turbine engine, excluding any nacelle; afirst spool including a first shaft connecting a first compressor andthe first turbine, the first shaft driving the propulsor through thegear train to drive the propulsor at a lower speed than the first spool,and the first compressor including three stages; and a second spoolincluding a second shaft connecting a second compressor and a secondturbine, the second compressor including at least as many stages as thefirst compressor, the second turbine includes two stages, the firstturbine including three turbine rotors, and wherein a weight of thepropulsor is greater than a weight of the first turbine.
 2. The gasturbine engine as recited in claim 1, wherein the gear reduction ratiois greater than 2.6.
 3. The gas turbine engine as recited in claim 2,wherein the array has no more than 20 blades extending from the hub. 4.The gas turbine engine as recited in claim 3, wherein the secondcompressor includes a greater number of stages than the firstcompressor.
 5. The gas turbine engine as recited in claim 4, furthercomprising a compressor section including the first compressor and thesecond compressor, wherein a axially forwardmost blade row of thecompressor section is axially aft of the epicyclic gear train relativeto an engine longitudinal axis.
 6. The gas turbine engine as recited inclaim 4, wherein the weight of the propulsor assembly is no more than33.2 percent of the total weight of the gas turbine engine, excludingany nacelle.
 7. The gas turbine engine as recited in claim 4, whereinthe first compressor has a greater number of stages than the secondturbine.
 8. The gas turbine engine as recited in claim 7, wherein aratio between a total number of the blades of the propulsor and a totalnumber of turbine rotors of the first turbine is between 3.3 and 8.6. 9.The gas turbine engine as recited in claim 8, wherein the weight of thepropulsor assembly is at least 28.4 percent of the total weight of thegas turbine engine, excluding the nacelle.
 10. The gas turbine engine asrecited in claim 9, wherein the epicyclic gear train is a star gearsystem.
 11. The gas turbine engine as recited in claim 9, wherein theepicyclic gear train is a planetary gear system.
 12. The gas turbineengine as recited in claim 9, wherein the gas turbine engine is atwo-spool engine, the first spool is a low speed spool, and the secondspool is a high speed spool.
 13. The gas turbine engine as recited inclaim 12, wherein the first turbine includes no more than six turbinerotors.
 14. The gas turbine engine as recited in claim 13, wherein thefirst turbine includes an inlet, an outlet, and a pressure ratio greaterthan 5:1, wherein the pressure ratio is a ratio of a pressure measuredprior to the inlet as related to a pressure at the outlet prior to anyexhaust nozzle.
 15. The gas turbine engine as recited in claim 14,further comprising a mid-turbine frame between the first turbine and thesecond turbine, wherein the mid-turbine frame supports a bearing systemand includes a plurality of vanes that serve as inlet guide vanes forthe first turbine.
 16. The gas turbine engine as recited in claim 15,further comprising a compressor section including the first compressorand the second compressor, wherein a axially forwardmost blade row ofthe compressor section is axially aft of the epicyclic gear trainrelative to an engine longitudinal axis.
 17. The gas turbine engine asrecited in claim 14, wherein the first turbine has an equal or greaternumber of stages than the first compressor.
 18. The gas turbine engineas recited in claim 17, further comprising a mid-turbine frame betweenthe first turbine and the second turbine, wherein the mid-turbine framesupports a bearing system and includes a plurality of vanes that serveas inlet guide vanes for the first turbine.
 19. The gas turbine engineas recited in claim 18, wherein the weight of the first turbine is atleast 321 pounds.
 20. The gas turbine engine as recited in claim 18,wherein the weight of the propulsor is at least 712 pounds.
 21. The gasturbine engine as recited in claim 18, wherein the first turbine is athree-stage or four-stage turbine.
 22. The gas turbine engine as recitedin claim 21, wherein the weight of the propulsor assembly is 28.4percent to 33.2 percent of the total weight of the gas turbine engine,excluding the nacelle.
 23. The gas turbine engine as recited in claim 9,wherein the propulsor is a fan, and further comprising a bypass ratio ofgreater than 10 and a nacelle circumscribing the fan, wherein thepropulsor assembly excludes the nacelle, and the total weight of the gasturbine engine excludes the nacelle.
 24. The gas turbine engine asrecited in claim 23, further comprising a low fan pressure ratio of lessthan 1.45 at cruise at 0.8 Mach and 35,000 feet, the fan pressure ratiomeasured across the blade alone.
 25. The gas turbine engine as recitedin claim 24, wherein the epicyclic gear train is a star gear system. 26.The gas turbine engine as recited in claim 24, wherein the epicyclicgear train is a planetary gear system.
 27. The gas turbine engine asrecited in claim 26, wherein the gas turbine engine is a two-spoolengine, the first spool is a low speed spool, and the second spool is ahigh speed spool.
 28. The gas turbine engine as recited in claim 27,wherein the weight of the propulsor assembly is 28.4 percent to 33.2percent of the total weight of the gas turbine engine, excluding anynacelle.
 29. The gas turbine engine as recited in claim 28, furthercomprising a mid-turbine frame between the first turbine and the secondturbine, wherein the mid-turbine frame supports a bearing system andincludes a plurality of vanes that serve as inlet guide vanes for thefirst turbine.
 30. The gas turbine engine as recited in claim 29,wherein the first turbine includes no more than six turbine rotors.